A typical gas turbine engine includes a fan, a compressor, a combustor, and a turbine. Fuel and compressed air discharged from the compressor are mixed and burned in the combustor. The resulting hot combustion gases (e.g., comprising products of combustion and unburned air) are directed through a conduit section from the combustor to the turbine where the gases expand, converting thermal energy into mechanical energy in the form of turbine shaft rotation. In electric power applications, the turbine shaft is coupled to turn a generator. Many parts of the combustor section and turbine section are exposed directly to the hot combustion gasses, including the combustor, the transition duct between the combustor and the turbine section, including the airfoils and surrounding ring segments.
With the efficiency of a gas turbine engine increasing with the firing temperature of the combustion gas, it is desirable to increase the temperature of the combustion gases. Temperature limitations of the materials with which the engine and turbine components are formed limit the operating temperatures. Special superalloy materials have been developed for use in such high temperature environments. However, modern high efficiency combustion turbines have firing temperatures in excess of 1,600 degrees C., which is well in excess of the safe operating temperature of the structural materials used in the hot gas flow path components. Consequently, specific cooling arrangements, including film cooling, backside cooling and insulation coatings are used to protect the integrity of such components under these high temperature conditions. Airfoils are exemplary. The term airfoil as used herein refers to a turbine airfoil which may be a rotor (rotatable) blade or a stator (stationary) vane.
Ceramic and ceramic matrix composite (CMC) materials offer the potential for higher operating temperatures than do metal alloy materials, due to the inherent refractory nature of ceramic materials. This capability may be translated into a reduced cooling requirement that, in turn, may result in higher power, greater efficiency, and/or reduced emissions from the engine. Although CMC's are beginning to find applications in military aircraft engines, they are not commonly utilized in large industrial gas turbines. Ceramics and CMCs do not exhibit the necessary balance of mechanical properties required for the manufacture of entire components. Consequently the use of ceramics has been mainly limited to their application as Thermal Barrier Coatings.
Turbine airfoils normally have associated shrouds or platforms. An airfoil platform defines a flow path between adjacent airfoil members for directing the hot combustion gases past the airfoil. The platform is exposed to the same high temperature gas environment as the ceramic airfoil and thus may be formed of a ceramic material. Currently the majority of blades and vanes are cast as single piece monolithic components. However, to support the introduction of more advanced materials (such as ceramics, CMCs, intermetallics and refractory alloys) modular designs are being considered. In many modular designs the platform and the airfoil are formed as separate components that are mechanically joined together, as illustrated in U.S. Pat. No. 5,226,789. Such mechanical joints must be robust. They tend to be complicated and expensive.
Monolithic ceramic is readily moldable to a form, but it is limited to small shapes and is insufficiently strain-tolerant for robust designs. CMC materials incorporate ceramic fibers in a ceramic matrix for enhanced mechanical strength. However, conventional ceramic composite processing methods increase in complexity and cost in a complex three-dimensional component such as a turbine vane. U.S. Pat. No. 6,200,092 describes a turbine nozzle assembly having a vane forward segment formed of CMC material wherein the reinforcing fibers are specially oriented across the juncture of the airfoil and the platform members. Such special fiber placement in the airfoil-to-platform transition region presents a manufacturing challenge, especially with insulated CMC construction. Furthermore, for some CMC compositions, shrinkage during processing may result in residual stresses in complex shapes that are geometrically constrained. The airfoil-to-platform attachment area is one area where such stresses can arise. Additionally, load transfer between the airfoil and the platform results in interlaminar stresses in the fillet region where mechanical properties may be compromised.
In one solution to these problems, U.S. Pat. No. 6,648,597 discloses a method of manufacture for a vane component of a gas turbine where both the airfoil member and the platform member are formed of a ceramic matrix composite material, and joint surfaces of the airfoil member and the platform member are bonded together. The method may be performed by urging the respective joint surfaces of the airfoil member and the platform member together while both members are in a green body state, followed by forming a sinter bond between them. The sinter bond method may be densified with a matrix infiltration process and the sinter bond may be reinforced with a fastener connected between the respective joint surfaces. The joint surfaces may be bonded with an adhesive. However, ceramic joints using refractory adhesives alone are weak and unreliable for carrying primary loads. Furthermore, when such adhesives are applied to already-fired CMC parts in constrained geometries, the adhesives shrink and produce bond joint cracking.
Like reference numerals refer to like parts throughout the drawings.